Airplane manufacturers are under increasing pressure to produce lightweight, strong, and durable aircraft at the lowest cost for manufacture and lifecycle maintenance. An airplane must have sufficient structural strength to withstand stresses during flight, while being as light as possible to maximize the performance of the airplane. To address these concerns, aircraft manufacturers have increasingly used fiber-reinforced resin matrix composites.
These composites provide improved strength, fatigue resistance, stiffness, and strength-to-weight ratio by incorporating strong, stiff, carbon fibers into a softer, more ductile resin matrix. The resin matrix material transmits forces to the fibers and provides ductility and toughness, while the fibers carry most of the applied force. Unidirectional continuous fibers can produce anisotropic properties, while woven fabrics produce quasi-isotropic properties. Honeycomb core is often sandwiched between composite sheets to provide stiff panels having the highest specific strength.
Because of the noise regulations governing commercial transport aircraft, high bypass engines incorporate acoustic panels within the nacelles. Conventionally, these elements are made with an inner perforated skin, a surrounding buried septum honeycomb core, and a non-perforated outer skin, such as described in U.S. Pat. Nos. 4,600,619; 4,421,201; 4,235,303; 4,257,998; and 4,265,955, which we incorporate by reference. The inner and outer skins are metal, usually aluminum, or composite, and the honeycomb core is either aluminum or composite. Manufacturing these acoustic panels is a challenge because of their size, their complex curvature and the close tolerances necessary for them to function properly.
As shown in FIG. 1, a nacelle 10 for a commercial high bypass jet engine includes a thrust reverser assembly having a fore-and-aft translating sleeve 11 to cover or expose thrust reverser cascades 12 when deploying thrust reverser blocker doors 15 carried on the translating sleeve. The thrust reverser assembly is positioned just aft of a jet engine, not shown, as is used on an airplane. The thrust reverser assembly is fitted within the nacelle 10. The thrust reverser cascades 12 are circumferentially spaced around the interior of the nacelle.
During normal flying operations the translating sleeve 11 is in a closed, or forward, position to cover the thrust reverser cascades 12. For landing an airplane, the translating sleeve 11 is moved from the closed position to the rearwardly extended, or deployed, position by means of actuator rods 18. This positioning routes exhaust gas to flow through the thrust reverser cascades 12 so as to slow down the aircraft on the ground. Exhaust is rerouted through the thrust reverser cascades 12 by closing the circumferentially positioned blocker doors 15.
The translating sleeve 11 is usually formed from a pair of semi-cylindrical outer cowl panels 13 (only one shown in FIG. 2) and a pair of semi-cylindrical inner acoustic panels 14 (only one shown in FIG. 2) bonded together to form the aft portion of the cylindrical nacelle 10. The outer cowl and acoustic panels 13, 14 are bonded at their aft ends and branch or diverge to provide a chamber for containing and concealing the thrust reverser cascades 12 and the associated support structures.
When the translating sleeve 11 is in the stowed position (FIG. 2), the leading ends of the acoustic panel 14 and the outer cowl panel 13 extend on opposite sides of the thrust reverser cascades 12. When the thrust reverser is deployed, the translating sleeve 11 is moved aft to expose the cascades 12 FIG. 3). The fan duct blocker doors 15 at the forward end of the acoustic panel 14 are deployed to divert fan flow through the cascades 12. The blocker door assembly is described in U.S. Pat. No. 4,852,805.
To form an acoustic composite sandwich panel, prior art methods used a male lay-up mandrel. The perforated composite inner skin was laid against the upper surface of the mandrel and buried septum honeycomb core was laid over the inner perforated skin. A composite non-perforated skin was then laid over the honeycomb core, and the three layers were cured or co-cured so as to form a single part.
This method did not provide index control for the inner or outer surface of the perforated sheets. Inexact tolerances on the inner and outer surfaces made locating and attaching details on the inside or outside of the acoustic panel difficult.
Thermal residual stresses produced during the curing process caused the acoustic panels to warp. Although the warpage was predictable to some extent, it was usually not uniform over the entire surface, leaving the part less than the desired design shape. Joining the acoustic panel and the outer cowl panel at a continuous aft joint with a smooth connection was difficult. Significant rework and shimming were required to correctly position the outer cowl panel and attach fittings against the outer side of the warped acoustic panel to complete the connection. Resin flowed into the perforations of the honeycomb core during the curing process, requiring rework of the perforated surface.
The acoustic panel must include recesses for receiving the fan duct blocker doors to provide a streamlined continuation during normal operation of the engine. In addition, deployment of the fan duct blocker doors imposes large bending moments at the leading end of the acoustic panel. As can be seen in FIG. 2, to receive the fan duct blocker doors 15 and to oppose the load of the fan duct blocker doors 15 when the thrust reverser is deployed, prior art acoustic panels 14 typically included a separate diaphragm 16 that was fastened at the leading end of the acoustic panel 14, and reinforced by gussets (not shown). An aft ring 17 extended between the acoustic panel 14 and the diaphragm 16. The blocker doors 15 were hinged from the leading end of the diaphragm 16. A forward ring 18, sometimes called a "bullnose ring", extended from the leading edge of the diaphragm 16 toward the thrust reverser cascades 12.
Because the gussets, the forward ring 18, the aft ring 17, and the diaphragm 16 were separate pieces, assembly of the thrust reverser and acoustic panel was laborious. The associated fasteners and connecting parts added significant weight to the acoustic panel 14. The steep angles formed by forward and aft rings deterred anyone from trying to form them in a single piece. Reducing the number of parts will reduce assembly time and will improve performance because an integral part made to close tolerance with the method of the present invention will be more aerodynamically efficient.
The fan duct blocker doors 15 fold downward and fit within recesses on the inner side of the acoustic panel. The blocker doors are trapezoidal so when stored, they create a triangular gap that needed to be filled for proper efficient airflow. The triangular gaps were filled by separate "wedge fairings" that were difficult to install with precision. The wedge fairings did not provide significant sound absorption. Attempts to form the wedge fairings integrally with the acoustic panel have not been successful.
The acoustic panels usually require reinforcement in the areas of attachment so that stresses applied by fittings attached to the acoustic panel will not damage the skins or core during sustained ultimate loads. To provide support at areas of fastened detail, prior art added plies to the skins to make the areas for fastening thicker than surrounding areas of the panel. The plies decreased in width so that the edges of the added plies formed "steps" or "ramps". The ramps help to dissipate forces applied through the fasteners into the skin. The presence of extra composite material at the ramps meant that perforations cannot be practically provided in the area covered by the ramps. Therefore, use of the ramps decreases the sound absorption area of the acoustic panel. There is a need for an acoustic panel that provides reinforcement in areas of fastened detail with a minimal loss of acoustic absorbing area.
Because the buried septum honeycomb core has little compressive and shear strength in directions parallel to the panel surface, it is often necessary to reinforce the honeycomb core in areas around fasteners and along the edges of the panels. Often, a dense core is substituted for the honeycomb core in the area to receive a fastener. Alternatively, portions of the honeycomb core can be removed and replaced by a potting compound. Potting compounds are also used along the edges of the panels. Each of these solutions poses problems. Dense cores are expensive and require additional processing steps to insert. Potting compounds are heavy, often disconnect with the composite skins, and require extensive labor to apply. There is a need for a more efficient way of providing support for a fastener in a composite. In addition, there is a need for a more efficient manner of providing solid edges ("closeouts") for a panel.
Composite panels are often tested prior to use by ultrasonic inspection. During a typical ultrasonic inspection, a Through Transmission Ultrasonic (TTU) sender is mounted on the opposite side of honeycomb-core composite panel from a TTU receiver. The TTU sender and the TTU receiver each have a water column that extends to the honeycomb-core composite panel. The TTU sender sends a signal that propagates through its water column, through the honeycomb-core composite panel, through the water column of the TTU receiver, and to the TTU receiver. Variations in the signal resonant frequency received by the TTU receiver indicate either changes in internal structure of the panel or internal flaws within the composite assembly.
A problem occurs when the TTU sender and TTU receiver approach an area of angular change in the honeycomb-core composite panel. If the water column is not extending parallel to the surface of the part, the signal path can be altered, causing inaccurate data from the TTU printout. There is a need for a method of more accurately performing nondestructive inspection in areas of angular change in a honeycomb-core composite panel.